Compressor section of gas turbine engine including hybrid shroud with casing treatment and abradable section

Patent metadata

Field	Value
Publication number	US-10876423-B2
Application number	US-201816235876-A
Country	US
Kind code	B2
Filing date	Dec 28, 2018
Priority date	Dec 28, 2018
Publication date	Dec 29, 2020
Grant date	Dec 29, 2020

Abstract

A gas turbine engine includes a shroud with an abradable section and a non-abradable section that cooperatively define a shroud surface. The gas turbine engine also includes a rotor that is supported for rotation within the shroud to generate an aft axial fluid flow. The rotor includes a blade with a blade tip that is crowned and that opposes the abradable section and the non-abradable section of the shroud surface. A crown area of the blade tip opposes the abradable section. A casing treatment feature is provided in the non-abradable section of the shroud to oppose the blade tip of the rotor.

First claim

What is claimed is:

1. A gas turbine engine comprising:
a shroud with an abradable section and a non-abradable section that cooperatively define a shroud surface;
a rotor that is supported for rotation about a longitudinal axis within the shroud to generate an aft axial fluid flow, the rotor including a blade with a blade tip that extends axially between a leading edge and a trailing edge of the blade, that is crowned and that opposes the abradable section and the non-abradable section of the shroud surface, a crown area of the blade tip opposing the abradable section; and
a casing treatment feature that is provided in the non-abradable section of the shroud to oppose the blade tip of the rotor;
wherein, in a projection of the blade tip onto a longitudinal plane, a theta angle is defined between an imaginary axial line and an imaginary tangential line, the imaginary axial line being parallel to the longitudinal axis, the imaginary tangential line being tangential to the blade tip; and
wherein a change in the theta angle along the blade tip in a downstream direction is, at most, zero.

2. The gas turbine engine of claim 1 ,
wherein a clearance region is defined between the blade tip and the shroud surface;
wherein a crown clearance dimension measured between the shroud surface and the blade tip at the crown area is less than a leading clearance dimension and a trailing clearance dimension, the leading clearance dimension measured between the shroud surface and the blade tip proximate the leading edge, the trailing clearance dimension measured between the shroud surface and the blade tip proximate the trailing edge.

3. The gas turbine engine of claim 2 , wherein the crown area clearance dimension is between approximately forty percent (40%) to sixty percent (60%) of the leading edge clearance dimension.

4. The gas turbine engine of claim 2 , wherein the blade tip has a radius that changes continuously from the leading edge to the trailing edge.

5. The gas turbine engine of claim 1 ,
wherein the shroud has a radius that remains substantially constant in a downstream direction relative to the longitudinal axis.

6. The gas turbine engine of claim 1 ,
wherein the shroud radially tapers in a downstream direction relative to the longitudinal axis.

7. The gas turbine engine of claim 1 , wherein the theta angle proximate the leading edge is a positive angle.

8. The gas turbine engine of claim 1 , wherein the theta angle proximate the leading edge is a negative angle.

9. The gas turbine engine of claim 1 , wherein the theta angle changes continuously along an entirety of the blade tip in the downstream direction.

10. The gas turbine engine of claim 1 , wherein the shroud includes a base material;
wherein the base material defines the non-abradable section of the shroud;
wherein the abradable section includes an upstream end and an inner diameter surface, the upstream end being embedded within the base material, and the inner diameter surface being exposed from the base material to partly define the shroud surface.

11. The gas turbine engine of claim 1 , wherein the casing treatment includes at least one of an aperture that is recessed into the shroud surface, a honeycomb structure that partly defines the shroud surface, a suction device, a blowing device, an active clearance control device, and a plasma flow control device.

12. The gas turbine engine of claim 1 , wherein the blade tip opposes the shroud surface to cooperatively define a clearance region therebetween, the clearance region having a flow axis;
wherein the abradable section includes an upstream end; and
wherein the crown area is disposed downstream of the upstream end relative to the flow axis.

13. A compressor section of a gas turbine engine comprising:
a shroud with an abradable section and a non-abradable section that cooperatively define a shroud surface;
a rotor that is supported for rotation about a longitudinal axis, the rotor including a blade with an blade tip that extends between a leading edge and a trailing edge of the blade, the blade tip opposing the abradable and non-abradable section of the shroud surface to define a clearance region between the blade tip and the shroud surface, a crown area of the blade tip opposing the abradable section;
a casing treatment feature that is recessed into the non-abradable section of the shroud surface to oppose the blade tip of the rotor;
wherein, in a projection of the blade tip onto a longitudinal plane, a theta angle is defined between an imaginary axial line and an imaginary tangential line, the imaginary axial line being parallel to the longitudinal axis, the imaginary tangential line being tangential to the blade tip; and
wherein a change in the theta angle along the blade tip in a downstream direction is, at most, zero.

14. The compressor section of claim 13 , wherein the theta angle proximate the leading edge is a positive angle.

15. The compressor section of claim 13 , wherein the theta angle proximate the leading edge is a negative angle.

16. The compressor section of claim 13 , wherein the theta angle changes continuously across an entirety of the blade tip in the downstream direction.

17. The compressor section of claim 13 , wherein the shroud includes a base material;
wherein the base material defines the non-abradable section of the shroud;
wherein the abradable section includes an upstream end, a downstream end, and an inner diameter surface, the upstream end and the downstream end being embedded within the base material, and the inner diameter surface being exposed from the base material to partly define the shroud surface.

18. A gas turbine engine comprising:
a shroud with an abradable section and a non-abradable section that cooperatively define a shroud surface;
a rotor that is supported for rotation within the shroud to generate an aft axial fluid flow, the rotor including a blade with a blade tip that extends axially between a leading edge and a trailing edge, that is crowned, and that opposes the abradable section and the non-abradable section of the shroud surface, a crown area of the blade tip opposing the abradable section, a clearance region being defined between the blade tip and the shroud surface;
a casing treatment feature that is provided in the non-abradable section of the shroud to oppose the blade tip of the rotor; and
a crown clearance dimension measured between the shroud surface and the blade tip at the crown area being less than a leading clearance dimension and a trailing clearance dimension, the leading clearance dimension measured between the shroud surface and the blade tip proximate the leading edge, the trailing clearance dimension measured between the shroud surface and the blade tip proximate the trailing edge, the crown area clearance dimension being between approximately forty percent (40%) to sixty percent (60%) of the leading edge clearance dimension.

19. The gas turbine engine of claim 18 , wherein the blade tip has a radius that changes continuously from the leading edge to the trailing edge.

20. The gas turbine engine of claim 18 , wherein the clearance region has a flow axis;
wherein the abradable section includes an upstream end; and
wherein the crown area is disposed downstream of the upstream end relative to the flow axis.

Assignees

• Honeywell Int Inc

Inventors

• Reynolds Bruce David
• Nolcheff Nick

Classifications

• F05D2240/307

  related to the tip of a rotor blade · CPC title

• F04D29/384

  characterised by form · CPC title

• F01D25/24

  Casings (modified for heating or cooling F01D25/14); Casing parts, e.g. diaphragms, casing fastenings (casings for rotary machines or engines in general F16M {; special arrangements in stators dealing with breaking-off of part of rotor F01D21/045}) · CPC title

• F01D5/141

  Shape, i.e. outer, aerodynamic form (F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147) · CPC title

• F05D2270/102

  caused by working fluid flow velocity profile distortion · CPC title